XFOIL Version 6.96 Calculated polar for: GOE 522 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.7054 0.10566 0.10243 -0.0904 0.9981 0.0284 -19.500 -0.7069 0.10064 0.09733 -0.0946 0.9973 0.0286 -19.250 -0.7073 0.09596 0.09256 -0.0986 0.9965 0.0286 -19.000 -0.7069 0.09147 0.08800 -0.1024 0.9955 0.0287 -18.750 -0.7037 0.08740 0.08384 -0.1060 0.9945 0.0288 -18.500 -0.6985 0.08363 0.08001 -0.1095 0.9935 0.0288 -18.250 -0.6918 0.08012 0.07643 -0.1128 0.9925 0.0288 -18.000 -0.6842 0.07679 0.07304 -0.1160 0.9916 0.0289 -17.750 -0.6812 0.07341 0.06961 -0.1184 0.9895 0.0289 -17.500 -0.6749 0.07024 0.06638 -0.1211 0.9877 0.0291 -17.250 -0.6654 0.06750 0.06360 -0.1235 0.9860 0.0291 -17.000 -0.6554 0.06469 0.06073 -0.1263 0.9844 0.0291 -16.750 -0.6426 0.06211 0.05811 -0.1290 0.9829 0.0291 -16.500 -0.6316 0.05913 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0.7010 0.01045 0.00426 -0.0887 0.4287 0.4185 2.750 0.7206 0.01060 0.00441 -0.0872 0.4244 0.4332 3.250 0.7660 0.01073 0.00466 -0.0853 0.4187 0.4678 3.500 0.7876 0.01082 0.00480 -0.0842 0.4152 0.4863 3.750 0.8083 0.01094 0.00495 -0.0830 0.4115 0.5043 4.000 0.8279 0.01111 0.00514 -0.0816 0.4075 0.5225 4.250 0.8476 0.01127 0.00533 -0.0801 0.4038 0.5416 4.500 0.8699 0.01135 0.00549 -0.0792 0.4010 0.5605 4.750 0.8911 0.01147 0.00566 -0.0781 0.3975 0.5785 5.000 0.9113 0.01163 0.00586 -0.0768 0.3935 0.5982 5.250 0.9296 0.01183 0.00609 -0.0753 0.3890 0.6189 5.500 0.9482 0.01202 0.00633 -0.0738 0.3851 0.6429 5.750 0.9687 0.01212 0.00654 -0.0726 0.3816 0.6719 6.000 0.9873 0.01226 0.00678 -0.0710 0.3773 0.7096 6.250 1.0016 0.01238 0.00709 -0.0687 0.3726 0.7752 6.500 1.1020 0.01246 0.00756 -0.0843 0.3645 0.9420 6.750 1.1478 0.01284 0.00793 -0.0885 0.3577 1.0000 7.000 1.1664 0.01327 0.00829 -0.0873 0.3509 1.0000 7.250 1.1889 0.01354 0.00856 -0.0867 0.3454 1.0000 7.500 1.2079 0.01397 0.00894 -0.0857 0.3379 1.0000 7.750 1.2264 0.01443 0.00936 -0.0846 0.3304 1.0000 8.000 1.2451 0.01490 0.00979 -0.0836 0.3214 1.0000 8.250 1.2622 0.01544 0.01029 -0.0824 0.3118 1.0000 8.500 1.2758 0.01617 0.01094 -0.0808 0.2995 1.0000 8.750 1.2911 0.01684 0.01154 -0.0794 0.2874 1.0000 9.000 1.3044 0.01762 0.01226 -0.0779 0.2757 1.0000 9.250 1.3170 0.01847 0.01305 -0.0763 0.2658 1.0000 9.500 1.3314 0.01924 0.01378 -0.0751 0.2577 1.0000 9.750 1.3440 0.02013 0.01463 -0.0736 0.2507 1.0000 10.000 1.3591 0.02089 0.01538 -0.0725 0.2451 1.0000 10.250 1.3716 0.02183 0.01629 -0.0712 0.2402 1.0000 10.500 1.3855 0.02272 0.01717 -0.0701 0.2358 1.0000 10.750 1.4019 0.02346 0.01792 -0.0693 0.2332 1.0000 11.000 1.4157 0.02438 0.01885 -0.0682 0.2303 1.0000 11.250 1.4289 0.02535 0.01982 -0.0672 0.2276 1.0000 11.500 1.4398 0.02649 0.02096 -0.0660 0.2248 1.0000 11.750 1.4506 0.02767 0.02215 -0.0648 0.2222 1.0000 12.000 1.4666 0.02851 0.02303 -0.0642 0.2213 1.0000 12.250 1.4812 0.02947 0.02402 -0.0634 0.2201 1.0000 12.500 1.4956 0.03044 0.02502 -0.0627 0.2186 1.0000 12.750 1.5083 0.03154 0.02615 -0.0619 0.2170 1.0000 13.000 1.5195 0.03279 0.02742 -0.0611 0.2153 1.0000 13.250 1.5304 0.03405 0.02871 -0.0602 0.2138 1.0000 13.500 1.5391 0.03553 0.03022 -0.0593 0.2124 1.0000 13.750 1.5452 0.03721 0.03190 -0.0582 0.2099 1.0000 14.000 1.5508 0.03897 0.03368 -0.0572 0.2083 1.0000 14.250 1.5632 0.04018 0.03495 -0.0566 0.2075 1.0000 14.500 1.5745 0.04153 0.03635 -0.0561 0.2068 1.0000 14.750 1.5859 0.04288 0.03774 -0.0556 0.2060 1.0000 15.000 1.5969 0.04426 0.03917 -0.0550 0.2051 1.0000 15.250 1.6058 0.04584 0.04081 -0.0545 0.2043 1.0000 15.500 1.6158 0.04732 0.04233 -0.0540 0.2032 1.0000 15.750 1.6233 0.04905 0.04410 -0.0534 0.2021 1.0000 16.000 1.6309 0.05079 0.04588 -0.0529 0.2009 1.0000 16.250 1.6373 0.05265 0.04779 -0.0524 0.1998 1.0000 16.500 1.6420 0.05470 0.04987 -0.0519 0.1986 1.0000 16.750 1.6468 0.05671 0.05191 -0.0514 0.1974 1.0000 17.000 1.6494 0.05892 0.05415 -0.0509 0.1956 1.0000 17.250 1.6557 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